The invention relates to aircraft environmental control systems powered by gas turbine engines for supplying aircraft system bleed air. In particular, the invention relates to an aircraft gas turbine engine powered bleed air supply system for an environmental control system wherein the excess pressure energy is returned to the engine to improve fuel efficiency and the system is also used to reduce drag on engine nacelles or other surfaces by pumping boundary layer air. Environmental control systems, commonly referred to as ECS systems, incorporate various pieces of equipment such as turbocompressors, regulating valves, heat exchangers, and other devices in a system which is often referred to as an ECS pack to condition engine bleed air. Modern day jet aircraft use turbocompressors in their environmental controls systems to condition bleed air for use in the cabin wherein the turbocompressors are powered by the same bleed that is conditioned for cabin refreshing air and which is usually supplied by the gas turbine engines which provide aircraft propulsion. Bleed air is conventionally taken from the engine compressor at a stage downstream of the variable vane compressor stages so as not to interfere with the operation of the variable vane stages which greatly enhance the efficiency of the gas turbine engine and greatly reduces the specific fuel consumption (SFC) of the engine.
The compressor bleed air is cooled by fan air in a heat exchanger and is then delivered to the environmental control system for controlling cabin air freshness, pressure, and temperature. The ECS conventionally includes two ECS packs mounted in ECS bays on different sides of the aircraft which receive compressor bleed air from the engines. The bleed after being used to power the ECS pack and refresh the cabin is then dumped overboard. All the energy remaining in the bleed air dumped overboard cost fuel and therefore represents a significant loss in specific fuel consumption.
Extraction of aircraft bleed air from the engine compressor has adverse affects on the propulsion cycle and engine life. Engine turbine power is needed to compress air and account for compressor inefficiency. Therefore, extra fuel consumption is always associated with gas turbine engine compressor bleed air (air which does not produce thrust). This extra fuel burned in the engine combustor results in higher gas temperature delivered to the engine turbine and reduction of turbine blade life. Such penalties must be incurred in order that the engine turbine provide extra power associated with bleed air.
It is not possible, without undue complexity, to always bleed the engine compressor stage which provides exactly the correct pressure needed for the aircraft anti-ice and ECS systems. Typically only two bleed ports ar provided. Therefore, the result is to bleed air which exceeds minimum pressure requirements resulting in even higher penalty to the engine cycle than would be required by the aircraft systems.
Most often the bleed air is not only at a higher than required pressure, it is also too hot. For reasons of fire safety, maximum bleed air temperature is usually limited to 350.degree. to 500.degree. F. Temperature control requires cooling the bleed air with a precooler. Most modern engines use fan air to cool compressor bleed air. Use of fan air imposes an additional penalty on fuel consumption. Further, the precooler is usually large and requires a fan air scoop which produces drag. A typical large turbofan engine will consume about 2% extra fuel and run at about 20.degree. F. hotter turbine temperature in order to provide aircraft system bleed air. The present invention addresses these problems and deficiencies characteristic of the prior art and conventional apparatus used to supply aircraft bleed air.
A second aspect of this invention concerns the engine air driven starter. Air starters are conventionally air powered turbines having planetary gearboxes which are usually connected to a high pressure rotor in driving relationship and mounted to the engine accessory gearbox. The starter turbine rotates at very high speed and drives the engine through a planetary gear system during engine acceleration to just below idle speed. Once the engine lights it begins to develop its own power and, at a speed below idle, accelerates away from the starter. An overrunning mechanical clutch allows the starter to disengage and then the starter air is shut off and the starter turbine comes to rest. During the remainder of the flight the starter is not used for any purpose and simply represents extra weight carried around by the aircraft.
Within a very narrow flight profile of the aircraft, the starter can sometimes be used for emergency engine relight, but only at conditions where the windmill speed of the engine is low enough that the starter clutch can be engaged without damage due to what is referred to as crash engagement. Engine starters can not be used during normal aircraft cruise conditions; where the only means for relight is from the freely windmilling engine. One advantage of the present invention is that it permits operation of the air starter during all aircraft flight conditions thereby avoiding the delay in engine relight which can be associated with flight conditions unfavorable for fast windmill relights. The present invention further enhances the solution to the relight problem by using the starter turbine during all operation conditions as a means for improving the performance of the auxiliary bleed air compressor.
A third advantage of this invention relates to aerodynamic drag associated with engine nacelles, wings, pylons, tail sections and other aircraft outer surfaces. As air flows on to and over a surface such as an engine nacelle and aircraft wing it progressively builds up a low velocity boundary layer of increasing thickness. Within this boundary layer a portion of the velocity component of free stream total pressure is converted to increased static pressure. As the result of rise in static pressure, boundary layer thickness, and diffusion a point is reached where back pressure causes an otherwise laminar boundary layer to become turbulent.
In the turbulent region, a considerable amount of total pressure is converted to static temperature represented thermodynamically as an increase in entropy. By the time the boundary layer leaves the surface, or in the particular case of an aircraft gas turbine engine the end of the nacelle, an unrecoverable loss in total pressure has occurred. The large entropy rise associated with turbulence is at the expense of air momentum. Turbulence also gives rise to increased static pressure which may increase the intensity of rearward acting pressure force on the surface. Now, if the boundary layer thickness is kept small, separation and turbulence will not occur or will be delayed and drag can be substantially reduced.
One way to avoid increase in boundary thickness is to pump or bleed off boundary layer air through holes in the surface. Boundary layer pumps or compressors would be desirable from an aerodynamic standpoint but, because of the relatively large air flow rates and added weight and complexity associated with effective boundary layer pumping or bleeding, the concept has not been adapted in modern aircraft and engines. Therefore in one embodiment of the invention, this invention provides a means for effectively and economically using the engine starting turbine to power a nacelle boundary layer bleed compressor to bleed off laminar flow boundary layer air from the nacelle to reduce drag. Yet another embodiment uses an ECS heat exchanger cooling fan to bleed off laminar flow boundary layer air from the wing to reduce drag.
A similar problem was addressed in and reference may be made to the patent application Ser. No. 07/489,150 entitled "AIRCRAFT ENGINE STARTER INTEGRATED BOUNDARY BLEED SYSTEM", invented by Samuel Davison, filed Mar. 6, 1990 and assigned to the sam assignee and incorporated herein by reference. Bleed air taken into the engine compressor also incurs a ram drag penalty (loss of momentum). Engine net thrust is equal to engine exhaust momentum minus inlet ram drag.
It is, therefore, an object of the present invention to provide a more efficient aircraft gas turbine engine by using the energy in the compressor bleed air that is conventionally wasted.
It is another object of the present invention to provide a lighter weight and more efficient and longer life aircraft gas turbine engine.
A further object of the present invention is to reduce or eliminate the need for engine fan supplied cooling air for the ECS bleed air precooler.
Another object of the present invention is to provide a fuel efficient system for supplying compressed air to the aircraft ECS systems.
Another object of the present invention is to provide the engine with a quick and reliable in flight restart or relight capability.
Yet another object of the present invention is to provide the engine with a starter that avoids the need for crash engagement for in flight relight.
A further object of the present invention is to reduce aircraft boundary layer drag in a fuel efficient manner.
Another object of the present invention is to us boundary layer bleed air to reduce base pressure drag in low pressure regions of the aircraft engine fan duct.
The objects and other features and advantages will become more readily apparent in the following description when taken in conjunction with the appended drawings.